1. Field of the Invention
The present invention relates generally to rotary kinetic fluid motors and pumps, and more specifically to turbine blade platforms and cooling thereof.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine produces a hot gas flow of extremely high temperature that passes through a turbine for converting the flow into mechanical power to drive the rotor shaft on which the turbine is connected. One method of increasing the efficiency of the engine is to increase the temperature of the hot gas flow into the turbine. However, the temperature is limited to the material properties of the parts within the turbine, specifically the first stage stationary vanes and rotor blades because these parts are exposed to the hottest gas temperatures.
One method used to allow for higher temperatures while using the same materials is to provide for cooling air to pass through the vanes and blades in order to prevent the parts from thermal damage. Since the cooling air must also be under high pressure in order to prevent backflow of the hot gas flow into the airfoils, the cooling air is generally bled off from the compressor at the appropriate stage. Thus, the engine efficiency is decreased because the work performed by the compressor on the bleed off air used for cooling is wasted. Therefore, another method for increasing the efficiency of the engine is to minimize the amount of cooling air used in the airfoils while maximizing the amount of cooling this air can provide. Elaborate and complex cooling passage designs has been proposed in order to attain this goal.
Turbine blades in a gas turbine engine generally include a root, an airfoil extending from the root, and a platform extending from the root to airfoil transition. The platform forms a flow path surface for the hot gas flow to prevent the flow from passing around the rotor shaft section. FIG. 1 shows a Prior Art technique used to cool the platform of a turbine blade. The blade 12 includes a pressure side platform 13 and a suction side platform 14, with a film cooling hole 16 in the platform 13 to supply cooling air from a dead rim cavity 17 formed below the two platforms. The blade is cooled by long length to diameter cooling channels and film cooling that yields successful results. U.S. Pat. No. 5,382,135 issued to Green on Jan. 17, 1995 entitled ROTOR BLADE WITH COOLED INTEGRAL PLATFORM discloses a blade with platform cooling of this design. This design utilizing film cooling for the entire blade platform requires cooling air supply pressure at the blade dead rim cavity 17 higher than the peak blade platform external gas side pressure. This induces high leakage flow around the blade attachment region and thus causes a performance penalty, reducing the engine efficiency.
Another method of cooling the blade platform is shown in FIG. 2. This design also utilizes long length-to-diameter cooling channels that are drilled from the platform edge to the airfoil cooling core. A pressure side long length-to-diameter cooling channel 26 is located on the pressure side platform 23 and connects with the cooling channel 25 in the blade. A suction side platform 24 includes a cooling channel 24 that also connects to the cooling channel 25. The dead rim cavity 27 is formed below the two adjacent platforms 23 and 24. In this design, the blade platform wall produces unacceptable stress levels at the airfoil cooling core and platform cooling channels interface locations, and thus yield a low blade life. This is primary due to the large mass at the front and back end of the blade attachment which constrains the blade platform expansion. In conjunction with the cooling channels are oriented at transverse to the primary direction of the stress field and high stress concentration associates with the cooling channels at the entrance location.
Studies have shown that over temperature occur at the blade platform pressure side location as well as the aft portion of the suction side platform edge and the aft section of the suction side platform versus the airfoil junction. To address this over temperature problem, the blade platform cooling design of FIG. 3 was proposed. Two long length-to-diameter cooling channels 31 parallel to the blade platform are used to cool the platform pressure side surface. Supply cooling hole 32 delivers cooling air to the channel 31, and an exit hole 33 discharges the cooling air. A large diameter cooling channel 34 running along the platform suction side edge is supplied with cooling air from a supply cooling hole 35 and discharges through an exit hole 36, with three small channels 37 branching off from the large channel 34 that cool the platform suction side surface. In general, the cooling air is fed into the long length-to-diameter channels 31 and 34 at the front end of the platform with the dead rim cavity air through the supply holes 32 and 35, and then discharged at the aft face of the blade platform through the exit holes 33 and 36.
It is an object of the present invention to provide for a cooling air design for a blade platform in a gas turbine engine that will be more efficient that the cited prior art designs, and therefore improve the efficiency of the gas turbine engine.